This invention relates generally to the design and development of spacecraft and, more particularly, to the design of unmanned spacecraft, such as earth-orbiting satellites. Traditionally, each spacecraft is considered unique to its intended purpose or mission. In general, a spacecraft includes a combination of complex subsystems, for performing such functions as propulsion, communications, power generation and storage, attitude control, and thermal control. Each subsystem has different requirements and specific functions, depending on the spacecraft mission. Even a simple spacecraft may have subsystems with hundreds of components that have to be uniquely interconnected and controlled to perform as required.
The process of spacecraft design, prior to the present invention, requires the designers to specify all of the complex interconnections among the subsystems and other components, and later to perform rigorous and costly integration and testing of all these components. Integration and testing at this high level of complexity is very manpower intensive, and typically benefits only a single spacecraft design. The high level of design complexity, coupled with the need to change the design for each mission, results in long, expensive, non-recurring development cycles, the cost of which can exceed by many times the recurring cost of individual spacecraft.
Some designers have introduced the concept of a standard spacecraft "bus," the intended purpose of which is to provide an "standard" vehicle for different space missions. However, such so-called standard buses often need many customizing modifications from one mission to the next. Even identical units used on the same spacecraft often require individual analysis, accommodation and documentation due to varying placement, orientation and accommodation within a spacecraft. Some initial progress has been made in reducing design cost by using standardized connector hardware and data buses for transmission of data between components or subsystems. Even with these standardized features, however, the spacecraft design process still requires an extremely costly design effort and a rigorous and detailed system integration and testing phase.
Spacecraft currently use external box structures to house electronics grouped by payload or bus functions. Typically, many such boxes are used on a single spacecraft, and a massive wiring harness interconnects the maze of boxes. Data buses are used only sparingly to reduce harness weight. Payloads are usually grouped on two or more thermally optimized panels with non-standard interconnections between panels, bus and antennas. Any change or addition of components normally disrupts the entire system-level design and often requires changes in the mechanical layout, the electrical interconnections or the thermal design. From one mission to the next, changes may have to be made in the payload throughput, propulsion system parameters, electrical power supply, wiring harness, thermal design or processor capabilities. Changes of this type greatly limit the amount of design reuse from one mission to the next, even when so-called standard spacecraft buses are used.
Ideally, a new spacecraft design process is needed, to provide shortened production schedules and increased reuse of existing designs, but without limiting mission flexibility. As will shortly become apparent from the description that follows, the present invention meets and exceeds these goals.